As is well known, in order to improve engine performance the aircraft engine industry has expended a great effort in its attempts to minimize the gap between the outer air seal or shroud and the tip of the turbine blades of aircraft gas turbine engines. For example, U.S. Pat. No. 4,069,662, granted to I. Redinger et al on Jan. 24, 1978 and assigned to United Technologies Corporation, the assignee of this patent application, discloses an active clearance control system that selectively impinges air on the engine's casing to shrink the casing and move the outer air seals closer to the tips of the turbine blades. Other systems have passively attempted to reduce the gap by flowing air at different temperature levels in proximity to the outer air seals to cause them to contract or expand.
Also well known is that the complexity of the problem is directly related to the use for which the engine will be put. For example, the maneuvers associated with fighter aircraft put demands on the aircraft engine that far outweigh those demanded by a commercial aircraft. The pilot of a fighter aircraft will select many more bodies, chops and transitory conditions than would a pilot of a commercial airliner. These conditions obviously impact the design of the engine which is particularly true of the turbine rotor's space relations to the outer air seal. These demands by the pilot cause heating and cooling of the turbine section structure, such that the shrinking and expansion and their rates affect the turbine parasitic leakage problem. Hence, any technical contribution that serves to reduce the gap while allowing the turbine to operate without undue rubbing the outer air seals is considered to be very important and significant since the leakage impacts the overall performance of the engine.
We have found that we can reduce the parasitic leakage by discretely discharging a portion of the cooling air internally of the turbine blades from the tip of the blade in a particular direction.